50 results about "Axial compressor" patented technology

The purpose of the present invention is to provide a multi-dimensional coupled design method for the through-flow aerodynamic layout of a gas turbine transition section-high pressure compressor. In the present invention, the transition section and the high-pressure compressor are regarded as a system for integrated coupled design, and full integration with the upstream transition section is achieved in all dimensions and links of the high-pressure compressor's aerodynamic design. This realizes the integrated and collaborative design of the through-flow layout of the transition section and the high-pressure compressor, enabling the aerodynamic design of the transition section-high pressure compressor system in different dimensions to enter a stage of systematization, parameterization and refinement. It effectively improves the aerodynamic performance of the high-pressure compressor, enhances the accuracy of aerodynamic design, saves a large amount of design iteration time, shortens the design cycle, and is highly suitable for engineering design applications. The present invention is not limited to the high-pressure compressor of gas turbines; it is also applicable to the aerodynamic design process of aero-engine high pressure compressors and various industrial axial compressors with transition through-flow structures.

The application discloses a high-pressure-ratio transonic axial flow compressor, which comprises a hub, front-row rotor blades, rear-row rotor blades and stator blades; the rear-row rotor blades are arranged between the front-row rotor blades and the stator blades; a plurality of front-row rotor blades and a plurality of rear-row rotor blades are uniformly arranged on the hub along a circumferential direction; and the front-row rotor blades and the rear-row rotor blades one-to-one correspond to form tandem rotors. The application realizes shock wave supercharging through the front-row rotor blades, and the rear-row rotor blades complete large airflow deflection; meanwhile, the application adopts an adsorption type stator to reduce or prevent airflow separation in a blade channel, thereby significantly improving a compressor stage pressure ratio and improving aerodynamic performance of the compressor. The application can realize a higher stage pressure ratio, and in the case of high inlet Mach number, high load and large airflow deflection angle, the application overcomes the problem that a high stage pressure ratio and efficiency are difficult to realize due to the fact that flow separation in the compressor cannot be controlled in the prior art.

Motor vehicle, with at least one air inlet (3) through which air can be introduced into a flow channel (6) of the motor vehicle when the motor vehicle is moving forward, and with at least one wind turbine (9) arranged in the flow channel (6), which has at least one wind wheel (12) driven by the air introduced into the flow channel (6) via the air inlet (3) and flowing through the flow channel (6), and a generator (13) driven by the wind wheel (12), by means of which electrical energy can be provided by driving the generator (13), characterized by: - at least one blower wheel (14) arranged in the flow channel (6) downstream of the wind turbine (12), by means of which the air flowing through the flow channel (6) is to be conveyed to at least one axial compressor (15) arranged downstream of the blower wheel (14) by driving the blower wheel (14), by means of which the air is to be compressed by driving the axial compressor (15), and - at least one additional electric machine (16) provided in addition to the generator (13), common to the blower wheel (14) and the axial compressor (15), by means of which the blower wheel (14) and the axial compressor (15) can be driven, and - a turbine (18) arranged outside the flow channel (6) and provided in addition to the wind turbine (9), - a branch line (19) by means of which a portion of the air introduced into the flow channel (6) via the air inlet (3) can be diverted from the flow channel (6) downstream of the wind turbine (12) and upstream of the blower wheel (14) and directed to the turbine (18), which can be driven by means of the diverted portion of the air, and - a second generator (20) arranged outside the flow channel (6), provided in addition to the generator (13) and in addition to the electric machine (16), and driven by the turbine (18), by means of which electrical energy can be provided by driving the second generator (20), and a return line (22) by means of which the diverted part of the air can be discharged from the turbine (18) and introduced into the flow channel (6), and wherein at least one Venturi nozzle (23) is arranged in the flow channel (6) downstream of the axial compressor (15), which can be flowed through at least by the air compressed by means of the axial compressor (15), and wherein the Venturi nozzle (23) can be supplied with the part of the air flowing through the return line (22) via the return line (22) and is thus permeable to that part of the air.

A moving blade (20) of an axial flow compressor has an airfoil part (22) of which the cross section forms an airfoil shape and which extends in a blade height direction perpendicular to the cross section. A leading edge part (23) including a leading edge (23l), a trailing edge part (24) including a trailing edge (24t), a negative pressure surface (25), and a positive pressure surface (26) are included. The shape of the outer edge on a virtual plane perpendicular to the blade height direction in the trailing edge part is an arc. Provided that the maximum blade height (W) is the maximum interval corresponding to a position in the blade height direction between the positive pressure surface and the negative pressure surface, the ratio of the arc radius (Rtc) of the trailing edge part to the maximum blade width (Wc) at the end on the tip side is greater than the ratio of the arc radius (Rtm) of the trailing edge part to the maximum blade width (Wm) at the midpoint in the blade height direction.

This invention discloses a rotor blade of a segmented stacked axial compressor and its design method. The design method includes: determining the axial position of the stacking line; determining the circumferential position of the stacking line: the stacking line is S-shaped in the circumferential direction, and the stacking line includes four control points A, B, C, and D sequentially from the blade root to the blade tip along the height direction; the coordinates of the centroid positions of each basic airfoil section in the first interval from point B to point C conform to a specific formula; and determining the shape of the stacking line based on the determined axial and circumferential positions. This invention, by setting the stacking line of the rotor blade to be S-shaped in the circumferential direction, generates a reverse pressure gradient in the radial direction, causing the gas at both ends to migrate towards the blade center, thereby improving the flow in the end region. This reduces the influence of secondary flow in the blade tip region; and by setting a specific function, the circumferential offset of the stacking line can be controlled by constructing a simple equation, improving the computational efficiency of optimizing the airfoil.

Axial compressor (10) with casing treatment for an aircraft turbomachine, comprising an annular casing (12) having at least one annular row of air outlet ports (20) and at least one annular row of air inlet ports (22), the air inlet ports (22) being connected to the air outlet ports (20) by channels (24) allowing air recirculation from the air inlet ports (22) to the air outlet ports (20), the compressor (10) further comprising a shell (32) which is housed in a slot (26) of the casing (12) and which has openings (34), this shell (32) being radially extendable and rotationally movable in the slot (26). Figure for the abbreviation: Figure 1

A multi-stage axial compressor comprises a rotating shaft on which a plurality of impeller blades are mounted, a casing surrounding the rotating shaft, the casing forming a flow path of a working fluid between the rotating shaft and the casing, a wall section having an annular shape and extending in a circumferential direction of the rotating shaft to surround the casing, the wall section forming a dispensing chamber having an annular shape and connected to the flow passage, a plurality of connection sections connected to the outer circumferential surface of the wall section, the connection sections forming respective outlet flow passages connected to the dispensing chamber, and a plurality of dispensing pipes connected to the respective connection sections.In a cross-sectional view orthogonal to the rotating shaft, of two corner regions where an inner surface of each of the connection sections and an inner surface of the wall section intersect or cross, at the corner region located on a rear side in a direction of rotation of the working fluid in the dispensing chamber, where θ1 is defined as an angle formed between an inner surface of the connection section and the inner surface of the wall section, the angle 01 is not greater than 225°.

The invention discloses an industrial axial flow compressor with a small aspect ratio, relates to the technical field of axial flow compressors, and solves the technical problems of low single-stage pressure ratio, low efficiency, insufficient reliability and the like of an existing industrial compressor. The industrial axial flow compressor comprises a welding type machine shell, a stationary blade inner casing, a stationary blade linkage adjusting mechanism, a rotor assembly, an inlet convergence casing, an outlet diffusion casing and other structures; the number of stages of the compressor is 6-12 according to actual working condition requirements, a controllable diffusion blade type with a small aspect ratio is adopted for blades, a combined flow channel with the equal outer diameter and the equal inner diameter is adopted, and angle proportion adjustment of inlet adjustable guide blades and 1-3 stages of stator blades can be achieved through a stator blade linkage adjusting mechanism.

This invention relates to the field of aerodynamic design technology for multi-stage axial compressors, and discloses a design method and system for passive stability enhancement of a multi-stage axial compressor through self-circulation. By calculating the aerodynamic performance of the multi-stage axial compressor and evaluating engine stability, the method identifies operating conditions where the multi-stage axial compressor has insufficient stability margin. Unsteady flow field simulations are then conducted under these conditions to determine the first stage stall. For the first stage rotor tip stall, a self-circulating casing is used for passive stability enhancement of the multi-stage compressor. The self-circulating casing structure is derived from sensitivity analysis of its design parameters, selecting the optimal design parameters to ensure an increase in the stability margin of the multi-stage axial compressor across the entire speed range. This satisfies the stability margin requirements of multi-stage axial compressors and avoids the problem of insufficient stability margin in high-load multi-stage axial compressors across the entire operating speed range.

To suppress the occurrence of surging in an axial flow compressor.SOLUTION: A moving blade of an axial flow compressor has an airfoil part which has a cross section forming an airfoil shape and extends in a blade height direction perpendicular to the cross section. The moving blade has a leading edge part including a leading edge, a trailing edge part including a trailing edge, a negative pressure surface, and a positive pressure surface. The shape of an outer edge on a virtual plane perpendicular to the blade height direction in the trailing edge part is an arc. Provided that the maximum blade height is the maximum interval corresponding to a position in the blade height direction between the positive pressure surface and the negative pressure surface, the ratio of the arc radius of the trailing edge part to the maximum blade width at the end on a tip side is greater than the ratio of the arc radius of the trailing edge part to the maximum blade width at the midpoint in the blade height direction.SELECTED DRAWING: Figure 5

A turbofan engine capable of reducing the air flow of an outer duct comprises multiple stages of fans, the outer duct with the reduced cross section area, an axial-flow air compressor, a combustion chamber, a turbine, an exhaust nozzle and the like, and the duct ratio does not exceed 1.0 or about 1.0; during operation after starting, the air flow of the outer duct with the cross section area reduced is reduced compared with the air flow of the outer duct with the cross section area unchanged, the air flow entering the core engine is increased due to the fact that the flow of the fan is unchanged, and the air flow is increased through operation of the engine. Compared with a conventional turbofan engine, the total supercharge ratio and the thrust-weight ratio of the turbofan engine are improved when the number of stages of a fan, an axial-flow compressor and a turbine, the fan diameter, the temperature in front of the turbine and the cross sectional area of an outer duct of the turbofan engine are not changed within the safe operation range of the mechanical strength and the high-temperature resistance of the turbofan engine, and compared with the conventional turbofan engine with the outer duct with the reduced cross sectional area, the bypass ratio of the outer duct is the same; however, the stages of the fan, the axial-flow compressor and the turbine are not changed, and the front temperature of the turbine is not changed.

The application relates to a supersonic rotor blade basic element airfoil design method, a rotor blade and an axial flow compressor, and belongs to the technical field of compressor blade design. Three sections of front, middle and rear are divided by adopting two shock wave surfaces in a rotor blade basic element channel, each section turning angle and relative Mach number are calculated, the rotor blade basic element airfoil is geometrically modeled based on the section turning angles, and the geometric modeling is verified and fine-tuned based on the relative Mach numbers, so that the problems of low efficiency and complex process of the supersonic blade basic element airfoil design method of the existing axial flow compressor are solved.

This invention belongs to the technical field of compressor aerodynamic stability analysis. It proposes a deterministic learning-based method and system for analyzing the aerodynamic stability of axial compressors. Using pulsating pressure data during the aerodynamic instability development process of an axial compressor, a deterministic learning identification algorithm based on radial basis function neural networks is used to model the nonlinear dynamics of the compressor's internal flow field, thereby obtaining dynamic diagrams of normal and instability precursor states. Dynamic indices are extracted from these diagrams to construct dynamic feature vectors describing the compressor's aerodynamic stability. Based on the dynamic feature vectors corresponding to different operating states, the dynamic classification boundary between normal and instability precursor states is determined. The distance between the dynamic feature vector of the compressor data under test and the dynamic classification boundary is calculated to determine the compressor's operating state. This invention enables early warning of aerodynamic instability and has good interpretability and engineering application value.

The invention aims to provide an adjustable guide vane electric direct drive mechanism of an axial flow compressor, which belongs to the field of axial flow compressors, and comprises an electric cylinder, a compressor and a pull and push dynamometer, a trunnion is arranged on the side of the electric cylinder, the trunnion is connected with an inclined lug seat, the inclined lug seat is fixed on a bracket, the bracket is fixed on a casing of the compressor, and the pull and push dynamometer is arranged on the bracket. One end of the pull and push dynamometer is connected with an extension rod of the electric cylinder through a double-thread screw, and the other end of the pull and push dynamometer is connected with the rotatable guide vane linkage ring through a connector. The telescopic rod of the electric cylinder is directly connected with the rotatable guide vane linkage ring through the Y-shaped connector, so that the middle switching link is reduced, the position precision of the rotatable guide vane during rotation is ensured, and the driving efficiency of the electric cylinder is improved; the S-shaped push-pull force meter is installed between the electric cylinder telescopic rod and the rotatable guide vane linkage ring, the maximum push-pull force needed by the gas compressor rotatable guide vane linkage ring can be displayed and tested in real time, the electric cylinder power level is reasonably selected, and unnecessary waste is avoided.

An end wall self-adaptive jet device of an axial flow compressor. It is installed on the outer surface of the casing corresponding to the position of the stator blade, which is a circular tapered pipe formed by sweeping along the center line, the center line is fitted by a Bezier curve, one end is an air inlet, the air inlet penetrates the casing behind the leading edge of the stator blade and closely adheres to the pressure surface of the stator blade, the other end is an air outlet, the air outlet penetrates the casing behind the middle of the axial chord length of the stator blade and is close to the suction surface of the stator blade. The effect of the application is that the pressure surface and the suction surface of the stator blade on the axial flow compressor are connected, the pressure difference between the air inlet and the air outlet forms a high-energy jet acting on the corner separation area, and the self-adaptive adjustment can be realized according to the change of the working condition of the axial flow compressor. Under suitable design parameters, it can effectively act on the corner flow field of the stator blade, reduce the range of corner separation, reduce the total pressure loss, improve the load of the stator blade, and improve the total pressure ratio and isentropic efficiency under small flow working condition.

Provided is an axial compressor comprising an exit guide blade that is provided on the downstream side of the final-stage stationary blade. When, in a cross section orthogonal to the blade height direction of the final-stage stationary blade, an extended straight line of a camber line of the final-stage stationary blade from the trailing edge of the final-stage stationary blade is referred to as a camber line extension line, the shortest distance between the camber line extension line and the negative pressure surface of the exit guide blade is Dmim, and the chord length of the exit guide blade is C, the final-stage stationary blade and the exit guide blade satisfy the expression 0.07 ≤ Dmin / C ≤ 0.20 in the entire range of the final-stage stationary blade in the blade height direction.

The invention belongs to the technical field of aero-engines, and particularly relates to a method for testing modal waves and surge waves of an axial-flow compressor, which comprises the following steps of: mounting a dynamic pressure test sensor matrix at the top section of a rotor blade of the axial-flow compressor; correcting the acquired sensor matrix data, and eliminating defective pixels and abnormal interference data; performing interpolation processing on the corrected data, including time interpolation along the rotation direction of the rotor and the chordwise direction of the blade, and space interpolation along the chordwise direction of the rotor blade; constructing a space-time grid by taking a chord-direction data point of the blade as a Y axis and taking a time sampling point and a rotor rotation direction sampling point as an X axis, and correcting a data arrangement sequence of the X axis by combining an included angle parameter between the blade and the axis, so that the grid corresponds to a space-time position of an actual measurement point; and mapping the interpolated data into the space-time grid to generate a dynamic pressure distribution cloud picture for analyzing the modal wave and the surge wave, so that the modal wave and the surge wave are visualized.

The utility model relates to the technical field of manufacturing of axial flow compressor blades, in particular to a self-adaptive positioning tool for remanufacturing of axial flow compressor blades, which comprises a positioning table, a first positioning table, a second positioning table, a first positioning table and a second positioning table. The clamp table is arranged above the positioning table, an open hole is formed in the clamp table, and the center of the open hole is coaxial with the center of the second fixing piece and coaxial with the center of the first fixing piece; and the multiple clamp assemblies are arranged on the clamp table in the circumferential direction, each clamp assembly comprises at least one curvature clamping block, and the curvature of the curvature clamping blocks is matched with the curvature of the molded surface of the target axial flow compressor blade. The top end and the molded surface of the target axial flow compressor blade are fixed based on the positioning table and the multiple clamp assemblies on the clamp table, the tensile stress and the torsional stress borne by the axial flow compressor blade due to machining are effectively relieved, and low-stress clamping and self-adaptive positioning in the remanufacturing process are achieved.

A gas turbine engine which includes an outer casing; a central longitudinal hollow shaft with a forward air inlet; a three stage rotating superconducting electric bypass fan with front and rear fan blades and a diffuser blade interposed between said front and rear fan blades wherein the diffuser blade rotates in an opposite direction to the front and rear fan blades; a multiple stage superconducting axial compressor positioned aft of the three stage rotating superconducting electric bypass fan; a multiple stage superconducting electric turbine core positioned aft of the multiple stage variable speed superconducting axial compressor, whereby the electric power from the multiple stage superconducting electric turbine core powers the three stage superconducting electric bypass fan and the multiple stage superconducting axial compressor.

The invention discloses an intelligent surge fault early warning method for a multi-stage high-pressure-ratio axial flow compressor of an aero-engine, and belongs to the technical field of aero-engine fault prediction. Firstly, pressure signals of the aero-engine in various operation states are collected to obtain original data, and a graph network is constructed; secondly, preprocessing the original data, and dividing the original data into a training set and a verification set; then, constructing a graph attention physical information neural network model composed of a graph attention layer, a probability prediction network and a physical information prediction network; training a graph attention physical information neural network model; then, verifying and finely adjusting the graph attention physical information neural network model; and finally, surge fault prediction is carried out on the multi-stage high-pressure-ratio axial flow compressor, and a prediction result is output to a surge early warning system. According to the method provided by the invention, the generalization ability of the model can be remarkably improved; an early warning signal can be given in advance, and a high-reliability solution is provided for surge prevention and control of the multi-stage high-pressure-ratio axial flow compressor of the aero-engine.

The present application is suitable for the compressor anti-surge control field, and provides a control method and device of axial flow compressor with inter-stage bleeding. In the embodiment, for the axial flow compressor with inter-stage bleeding, a start-stop control line group and a continuous operation control line group are set in advance according to the surge boundary and surge margin of the compressor. When the compressor is running, the matched control line group is determined according to the operation parameters, running state and switch state of the inter-stage bleeding valve of the compressor. Then, the valves of the compressor, such as the exhaust air supply valve, the exhaust air vent valve and the inter-stage bleeding valve, are adjusted according to the pressure ratio, equivalent speed and inlet guide vane opening of the compressor, so as to avoid the unstable operation of the compressor and the problems of possible stall or surge.

The object of the present invention is to provide a multidimensional coupling design method for the pneumatic flow arrangement of a transition section high-pressure compressor of a gas turbine. In the present invention, the transition section with the high-pressure compressor is designed as a system for a holistic coupling design.In all dimensions and segments of the pneumatic design of the high-pressure compressor, complete integration with the upstream transition section is achieved, and the integrated and synergistic design of the flow arrangement of the transition section and the high-pressure compressor is realized. This allows the pneumatic design of the transition section-high-pressure compressor in various dimensions to enter a systematic, parametric, and refined stage, effectively improving the pneumatic performance of the high-pressure compressor, increasing pneumatic design accuracy, saving considerable time in design iterations, shortening the design cycle, and making it highly suitable for technical design applications.The present invention is not limited to high-pressure compressors of a gas turbine, but is also applicable to high-pressure compressors for aerospace engines and various industrial axial compressors with transition flow structure in the pneumatic design process.

The invention discloses a multistage counter-rotating axial flow compressor order reduction design method suitable for gas-liquid two-phase mixed transportation. The multistage counter-rotating axial flow compressor order reduction design method comprises the following steps that a crossed double-shaft single-stage counter-rotating blade structure is designed; carrying out order reduction treatment on design parameters of the single-stage contra-rotating blade; carrying out order reduction treatment on design parameters of the multi-stage contra-rotating blade; carrying out order reduction optimization design on the multi-stage contra-rotating blades; and carrying out correction design on a gas-liquid mixed transportation flow field. The design method provided by the invention can provide an extremely important reference basis for deep sea environment pressurization design, and has extremely wide application prospects and extremely important practical significance.

Axial compressor (10) with casing treatment for an aircraft turbomachine, comprising an annular casing (12) having at least one annular row of air outlet ports (20) and at least one annular row of air inlet ports (22), the air inlet ports (22) being connected to the air outlet ports (20) by channels (24) allowing air recirculation from the air inlet ports (22) to the air outlet ports (20), the compressor (10) further comprising a shell (32) which is housed in a slot (26) of the casing (12) and which has openings (34), this shell (32) being radially extendable and rotationally movable in the slot (26). Figure for the abbreviation: Figure 1

The invention discloses a single-shaft turbojet engine. Comprising an air inlet necking used for initial acceleration of air flow, a fan provided with a blade profile suitable for the supersonic speed of the air flow, a starting-power generator arranged in the center of a fan wheel and based on a brushless direct current motor, a multi-stage axial flow air compressor used for compressing the air flow of a core machine, and an annular combustion chamber. And a turbine having active blades and having one or more wheels. The fan, the starter-generator, the axial flow compressor and the turbine are all mounted on a single shaft without reduction gears. Parameters of the blades, the fan, the compressor and the turbine are optimized, and the gear train does not need to be decelerated. The technical scheme can be used for engines with the bypass ratio as high as 8. According to the invention, the characteristics of weight and size of the turbojet engine can be significantly improved while good efficiency is maintained.

The present invention is directed to axial-centrifugal contra-rotating compressor system comprising combination of contra-rotating axial compressor stage with a back end centrifugal compressor stage, eliminating stator blades, an inter-stage connecting duct at the entry of the centrifugal compressor and an discharge diffuser at the exit of centrifugal compressor rotor characterized by radial to axial turning annular passage as per the gas turbine engine requirement, favouring reduction in engine size, weight, number of parts, overall engine length and produces a higher-pressure rise and swirl-free discharge towards the combustion chamber. The first low-pressure rotor having at least one axial blading assembly rotates in the direction opposite to the second high-pressure rotor blading assembly comprising of at least one axial rotor and one centrifugal rotor. The blades for axial compressor are adopted of low-aspect ratio configuration for reliable high load stall free operation. The inter-stage connecting duct minimizes total pressure loss. The discharge diffuser recovers higher exit velocity and hence higher kinetic energy.

The utility model relates to the technical field of axial flow compressors, in particular to a high-pressure balance pipeline system for an axial flow compressor. Comprising a high-pressure balance pipeline, one end of the high-pressure balance pipeline is connected with an exhaust high-pressure side of an axial flow compressor shell through an exhaust flange, the other end of the high-pressure balance pipeline is connected with an air inlet low-pressure side of the axial flow compressor shell through an air inlet flange, and a valve is arranged on the high-pressure balance pipeline; the temperature sensor is arranged on a main thrust surface and an auxiliary thrust surface of a thrust bearing of the axial flow compressor and is used for detecting the temperature of the main thrust surface and the auxiliary thrust surface of the thrust bearing of the axial flow compressor; the controller is connected with the valve and the temperature sensor and used for regulating and controlling the opening degree of the valve according to the detection result of the temperature sensor. According to the high-pressure balance pipeline system for the axial flow compressor, it is guaranteed that the thrust bearing of the axial flow compressor is in a good operation state, the load of the thrust bearing is reduced, and the service life of the thrust bearing is prolonged.

The application discloses an intelligent fault-tolerant control method of an axial flow compressor control system, which comprises the following steps: collecting throat differential pressure delta P, exhaust pressure P2 and inlet temperature T1 by adopting three-redundancy design; setting the local maximum temperature as the upper limit of temperature compensation; designing temperature sensor disconnection and automatic correction; and predicting the surge occurrence of the unit according to the change of the exhaust pressure P2 speed vector, and making an intervention action in advance. The intelligent fault-tolerant design method is used to automatically judge, compensate, inhibit, eliminate, correct faults in a combination of hardware fault tolerance and software fault tolerance, so as to ensure that the equipment continues to operate safely, efficiently and reliably, or to ensure that the unit completes the predetermined function within the specified time at the cost of part of the performance of the unit.

An axial compressor rotor assembly includes a plurality of rotor disks extending axially along and circumferentially about a central axis between an upstream end and a downstream end. The rotor disks extend radially between an inner radial hub and an outer radial rim, and axially between a forward rim end and a rearward rim end. An airfoil extends spanwise from each of the rotor disks. A first rotor disk is disposed forward of a second rotor disk. The second rotor disk includes an interstage shaft. The interstage shaft includes an outer radial end, an inner radial end, and an arm. The arm extending axially from an outer surface of the interstage shaft to a distal end. The rearward rim end of the first rotor disk is fixedly joined to the distal end of the arm, for example, by inertia welding.

To provide a control apparatus and a control method of variable stator blade capable of preventing the deterioration of service life of a gas-turbine engine by preventing the thrust of a driving device of the variable stator from rising.SOLUTION: A control apparatus of variable stator blade includes an arm 21 connected with a stationary vane 11 of an axial compressor 1, a rotary ring 22 which is connected with one end part of the arm 21 and is located on an outer periphery of a casing 10 of the axial compressor 1, a driving device 40 which rotates the stationary vane 11 via the arm 21 by rotating the rotary ring 22 and a controller CL which controls the driving device 40 by means of a command signal. The controller CL has a deceleration dead zone setting part 51 which performs deceleration dead zone control for invalidating the command signal during a predetermined time after displacing the stationary vane 11 to an objective installation angle θ, and an offset part 52 which returns the driving device 40 as many as a predetermined distance in such a direction as to relax an elastic energy stored in the arm 21, the rotary ring 22 and the driving device 40.SELECTED DRAWING: Figure 3